The invention relates to the field of attitude reference determination of moving vehicles. More particularly, the present invention relates to attitude reference determination of spacecraft providing carrier integer cycle ambiguity resolution.
Aircraft and spacecraft vehicles require methods and apparatus for accurately determining respective positions during flight missions. Position determination has been improved using the Global Position System (GPS) that includes a constellation of orbiting GPS satellites broadcasting ephemeral signals to receivers. The receivers may be fixed at a ground base or carried onboard a moving ground, airborne or space vehicle. Position determinations resulting in determined points in space as well as an attitude reference using GPS is well known in the art. During GPS position determination applications, a range distance between a receiver and a GPS satellite is determined by measuring the time it takes for the pseudo random signal to travel the distance from the GPS satellite to the receiver. Knowing four range measurements from four respective different GPS satellites to the receiver, the receiver position can be uniquely determined by well known spatial reference frame computational processes. The receiver position is the location of the receiver in three dimensional X, Y, Z space. The attitude of a vehicle is an angular orientation that requires the position of at least three receiver antennas to achieve centimeter level accuracy. However, the accuracy of these range measurements is of the order of several meters and therefore the range measurements are not precise enough for determining the attitude of the vehicle as an angular orientation reference. Another measurable quantity of the GPS signal is the carrier phase that is a fractional part of a carrier cycle. The carrier phase needs to be precisely determined for improved accuracy of the position determination. Before the fractional carrier phase measurements at the antennas can be converted into receiver to satellite ranges, knowledge of the integer cycles spanning these ranges is required. The integer cycles can be ambiguously determined due to a lack of precise range determinations. A number of approaches currently exist to resolve this integer cycle ambiguity. However, the current methods suffer various computational problems limiting the ability to effectively determine the carrier cycles. Earlier software based approaches utilize an integer search method attempting to minimize a cost function. The software based approaches check different sets of integer values that can number in millions for antennas separated by just a few meters. The problems that arise with this software based approach include the existence of nonunique solutions that create the possibility of converging to a wrong set of integers. Also, due to existence of a very large set of integer combinations, processing takes large amount of time to check all of the integer possibilities. This limits the reaction time of an agile vehicle. Further, most search algorithms may not always converge to a solution. Another current approach utilizes integer resolution algorithms that use additional information due to motion of the vehicle to determine the integer cycle of the carrier. These integer resolution algorithms possess the same inherent computational problems that are not significantly reduced. These and other disadvantages are solved or reduced using the invention.
An object of the invention is to provide a method for attitude reference determination.
Another object of the invention is to provide a method for determining elevation angles and azimuth angles to pseudo stars. Another object of the invention is to provide a method for determining elevation angles and azimuth angles to pseudo stars for attitude reference determination.
Yet another object of the invention is to provide a method for determining elevation and azimuth angles to GPS satellites by rotating receiver antennas about a reference axis and dithering the antennas along an orthogonal axes for carrier phase alignment of the received GPS signals.
Still another object of the invention is to provide a method for positioning receiver antennas in carrier phase alignment of GPS signals by rotation about a reference axis and dithering the receiver antennas along orthogonal axes for determining a position and attitude in inertial space.
A further object of the invention is to provide a method for positioning receiver antennas by rotation about an attitude axis and dithering along orthogonal axes for receiving GPS signals in carrier phase alignment when the antennas are orthogonally positioned for determining the attitude axis.
The invention is a method directed to positioning receiver antennas in carrier phase alignment of GPS signals for determining coelevation and azimuth angles to GPS satellites. The coelevation and azimuth angles can be used for determining the attitude of the vehicle in an inertial reference frame from known GPS satellite lines of sight. The method can be preferably used for determining the attitude of a airborne or spaceborne craft receiving GPS signals from GPS satellites functioning as pseudo star references. The method is practiced using a GPS receiver system and signal processing algorithms that determine the attitude of an arbitrary ground receiver, aircraft or spacecraft. In the preferred form,,the system uses a master antenna and two dependent slave antennas, each of which having a respective receiver for receiving the GPS signals. The antennas are orthogonally aligned respecting each other and are controlled to undergo prescribed motions relative to each other. A fractional phase of the GPS carrier signal received at each of two dependent slave antennas is measured relative to the master antenna. Processing of the measured carrier signal is used to eliminate the integer cycle ambiguity for determining the attitude reference by computing two noncolinear lines of sight vectors. The method comprises of means to determine more than one noncolinear unit vector along the lines of sight from the master antenna to GPS satellites functioning as pseudo stars.
The method includes signal processing steps to compute the unit vectors in two coordinate systems. The two coordinate systems are an earth centered inertial coordinate system and a local coordinate frame system attached to the craft. The signal processing steps are provided to compute the attitude from the direction cosine matrix between the local coordinate system and the earth centered inertial coordinate system. The system preferably includes the set of three antennas mounted on a rigid frame forming essentially two orthogonal directions. A first orthogonal direction extends from the first master antenna to the second x axis slave antenna along an x axis. A second orthogonal direction extends from the master antenna to a third y axis antenna along a y axis. The system provides motors for rotating the rigid frame about a vertical z axis extending through the master antenna. Each of the second and third slave antennas are controlled to undergo back-and-forth dither motion along respective orthogonal lines respectively extending from the master antenna to the second and third antennas. A controller rotates the frame about the z axis for positioning the rigid frame while controlling the two slave antennas to respectively undergo the dither motion along the x axis and y axis. A system includes a microprocessor that continuously measures the differential carrier phase between the master antenna and the x axis antenna and between the master antenna and y axis antenna. Signal processing steps of the differential phase data enables determination of the unit vectors along the line of sights from the master antenna and various GPS satellites. The lines of sight data is used for accurately determining the attitude reference of the vehicle. The position and attitude can be determined without resolving carrier phase ambiguity. These and other advantages will become more apparent from the following detailed description of the preferred embodiment.